1. Field of the invention
The present invention relates to a turbojet aero-engine with a high bypass ratio.
2. Description of the Prior Art
It has been common for some years to use bypass turbojet engines for subsonic propulsion, the advantage of mixing the cold and hot air streams being to increase the thrust of the hot engine.
Nevertheless, in practice, one is restricted as to the by-pass ratios which are possible to employ in such engines because of their frontal area and the large cross section which is the corollary of this. Above a bypass ratio of between 8 and 14, the engines become prohibitively large and are effectively prevented from being mounted below the wings of an aircraft.
If it is desired to exceed a bypass ratio of 8, the solution adopted in conventional turbojet engines with an upstream fan, at least up to a bypass ratio of about 11, is to provide the low pressure turbine with between 5 and 8 stages.
For a bypass ratio of from 11 to 14, the conventional solution is no longer convenient (owing to the need for an excesive number of low pressure turbine stages) and it is necessary to use a reduction gear for the front-mounted fan, or a rear-mounted contrarotating fan driven directly by a contrarotating power turbine with interleaved stages.
Above a bypass ratio of 14, one enters the realm of turbo-jet engines with high speed propellers.
These engines are of interest because of the improvement in specific fuel consumption which they provide, but have the disadvantage, because of their large overall diameter, of being capable of installation only at the rear and on either side of the fuselage of the aircraft, or only with highly-integrated and unconventional underwing installations of the type proposed in, for example, French Patent No. 2 622 507.